The present invention relates to a gas turbine engine and a gas turbine engine component. One aspect of the invention relates to the gas turbine engine component presenting an inner passage surface adapted to be surrounded by and to form a delimitation of a gas flow passage. Another aspect of the invention relates to the gas turbine engine component being adapted to provide a gas flow passage for at least a part of a gas flow through the gas turbine engine, which gas flow passage surrounds an inner passage surface, an outer passage surface forming an outer delimitation of said gas flow passage, a plurality of circumferentially spaced blades extending between the inner and outer passage surfaces.
In a gas turbine engine there are gas flow passages or ducts formed between an inner part and an outer part, surrounding the inner part. Often blades, in the form of struts or vanes are located in such passages, at which they are circumferentially distributed and extend between the inner and outer parts. The design of such passages is important for reasons of aerodynamic efficiency of the engine. For example, in known art, a lot of studies have been carried out concerning the design of end walls, herein also referred to as inner and outer passage surfaces, presented by such inner and outer parts, facing towards the gas flow passage, and at least partly extending between blades or vanes in turn extending between the inner and outer part. For example, testing of a cascade with an end, wall with a concave curvature near the blade suction surface is disclosed in “Non-axisymmetric End turbine end wall profiling”, Gregory-Smith D. G., Ingram G., Jayaraman, P., Harvey N. W., Rose M. G., Proceedings of the 1 MECH E Part A, Journal of Power and Energy, Vol. 215, Number 6, 2 Dec. 2001, pp. 721-734.
While studies of known art are focused on aerodynamic qualities of the passages, often the solutions provided are less than optimal regarding other requirements, such as manufacturing requirements, and structural characteristic requirements.
It is desirable to make the manufacturing of gas turbine engines easier.
According to an aspect of the invention, a gas turbine engine component is provided comprising an inner passage surface adapted to be surrounded by and to form a delimitation of a gas flow passage, wherein the trace of the inner passage surface in a cross-section perpendicular to a central axis of the component presents at least one substantially straight portion.
The central axis of the gas turbine engine component would normally be such that in a mounted state of the component in an engine, it coincides with an axis of rotation of at least one rotatable portion for the operation of a turbine assembly and a compressor assembly, located, in relation to a gas flow through the engine, downstream and upstream, respectively, of a combustion chamber.
Here, the trace portion of the inner passage surface being substantially straight means that it is straight, or that it deviates slightly from being straight in a manner that can be expected in practical use, for example due to deformation caused by manufacturing tolerances or thermal stresses.
The straight portion(s) of the trace of the inner passage surface provides, where the part presenting the inner passage surface is not divided into sections, better properties for thermal expansion, and, where the part presenting the inner passage surface is divided into sections, a simplified geometry for the joints between the sections, which in turn simplifies sealing between the inner part sections. Thus, the invention simplifies manufacturing of gas turbine engines.
Round shapes used in known solutions favour thermal stresses. The straight trace portion(s) of the inner passage surface provides for reduced risks of high thermal stresses. Thereby the structural performance of the part presenting the inner passage surface is increased. Such increase in the structural performance in turn provides for increasing the life of this part.
Preferably, the inner passage surface presents, in said cross-section, at least two substantially straight trace portions which have different extension directions. In yet more preferred embodiments, the inner passage surface has a polygonal shape. Thereby, the polygonal shape will be formed at least in said cross-section as a result of trace of the inner passage surface in the cross-section presenting a plurality of substantially straight portions.
Preferably, a plurality of circumferentially spaced blades extend from the inner passage surface. Preferably, at least one of the blades is located at an end point of a straight trace portion. In yet more preferred embodiments, at least one of the blades is located at a point of intersection between two adjacent substantially straight trace portions. Here, a blade being located at an end point of a straight trace portion means that it can be located on, close to, or in the vicinity of the end point. Similarly, a blade being located at a point of intersection between two adjacent substantially straight trace portions means that it can be located on, close to, or in the vicinity of the point of intersection.
Thereby gas flow passage portions between adjacent blades will, compared to traditional designs where the inner passage surface forms, in a cross-section perpendicular to said axis of rotation, a circular shape, have its cross-sectional area distributed more towards the region half-way between the blades. This will change the pressure distribution in the gas flow passage so as to reduce the risk of flow separation against delimitations of the gas flow passage. In other words, the gas flow passage cross-sectional area will be more concentrated to a central region between blades, which will change the pressure distribution in an advantageous manner so as to reduce secondary flow losses, especially at the blade roots. Thus, a gas turbine engine with improved aerodynamic qualities for gas flows, particularly for internal gas flows, will be provided.
It should be noted that the blades can be any sort in blades in the gas flow of the engine, for example turbine or compressor rotor blades, turbine or compressor stator vanes, or structural struts. More generally, the blades can be load-carrying, for example as struts, or not load-carrying, and the invention is applicable regardless whether the blades are adapted to redirect the gas flow.
In addition, the concentration of the passage cross-sectional area to a central region between blades will allow for the blade roots to be positioned further away from the axis of rotation of the engine. This will provide more space in an inside compartment, for example for oil conduits or air conduits, and it will also make assembly of the engine easier.
As further exemplified below, preferably a first and second angle between respective straight portions, at said cross-section, of the inner passage surface, on either side of the respective blade, and a straight line parallel to the spanwise direction of the respective blade, are both obtuse. As a result, the risk of flow separation at a region of intersection between the respective blade and the inner passage surface will be reduced. More specifically, in traditional designs, for example where the inner passage surface forms, in a cross-section perpendicular to said axis of rotation, a circular shape, said angles between the blades and the inner passage surface will be essentially right angles, or alternatively, one of them will be sharp. Such right and sharp angles will contribute to the stimulation of flow separation between the blade portion of maximum thickness and the blade trailing edge, where the gas flow decelerates. Opening said angles between the blades and the inner passage surface so as to be obtuse, reduces the risk of flow separation at the region of intersection between the respective blade and the inner passage surface.
As also mentioned below, the decreased risk of separation provides a special advantage where the gas flow passage is designed to provide diffusion, i.e. a decrease in the gas velocity, in that the passage can be designed so as to be shorter, contributing to reducing the length and weight of the engine.
Preferably, where at least one of the blades is located at an end point of a straight trace portion, and/or where at least one of the blades is located at a point of intersection between two adjacent substantially straight trace portions, the at least one straight trace portion of the inner passage surface is provided at all cross-sections, perpendicular to said central axis, along the entire blade in the direction of said central axis. This will further secure the aerodynamic benefits mentioned above.
Preferably, the engine comprises an inner part and an outer part forming between them the gas flow passage, which gas flow passage surrounds the inner part, the inner part presenting the inner passage surface. Thereby, the gas flow passage can present a generally circular shape as seen in a cross-section perpendicular to said axis of rotation. Also, thereby, where the circumferentially distributed blades are provided, they each extend between the inner and outer part.
In preferred embodiments, the inner passage surface presents in said cross-section a plurality of crests, and said at least one substantially straight portion of the trace is positioned between two adjacent crests. As exemplified below, said crests can form ridges extending at least partly in the direction of said axis of rotation, and blades can be located at the crests, i.e. on or in the vicinity of the crests. It should be noted that at intersections of adjacent straight trace portions, only one crest might be provided. However, for example where a relatively small concavity is provided at the intersection, which concavity extends mainly in the gas flow direction, two crests can be provided, each forming a ridge extending on a respective side of the concavity.
As exemplified below with reference to FIG. 7, preferably, a major part of the trace, at said cross-section, of the inner passage surface between any of two adjacent crests is substantially straight.
As exemplified below with reference to FIG. 8, preferably, ridges, forming said crests at said cross-section, are oriented in an angle to a plane being parallel with said axis of rotation. Preferably, said ridges are curved so that the angle of the local orientation of the ridges varies in the direction which is parallel to the axis of rotation.
In one embodiment, the inner passage surface forms an external surface of a hub for a propeller for a turboprop engine, or a hub for external blades for an unducted fan (UDF) or an open rotor.
According to another aspect of the invention, a gas turbine engine component is adapted to provide a gas flow passage for at least a part of a gas flow through the gas turbine engine, which gas flow passage surrounds an inner passage surface, an outer passage surface forming an outer delimitation of said gas flow passage, a plurality of circumferentially spaced blades extending between the inner and outer passage surfaces, wherein two portions of the trace, in a cross-section perpendicular to a central axis of the component, of the outer passage surface between two adjacent blades are substantially straight and oriented in an angle in relation to each other so as to form a concavity in the gas flow passage.
As exemplified below, the gas flow passage can be formed by an inner part and an outer part forming between them the gas flow passage and presenting the inner and outer passage surfaces, respectively. Also, the gas flow passage can surround the inner part so as to present a generally circular shape as seen in a cross-section perpendicular to said axis of rotation.
The advantages at manufacturing, mentioned above, are provided also at this aspect of the invention; i.e. the straight portions of the outer passage surface provide, where the part presenting the outer passage surface is not divided into sections, better properties for thermal expansion, and, where the part presenting the outer passage surface is divided into sections, a simplified geometry for the joints between the sections, which in turn simplifies sealing between the outer part sections.
Also, similarly to embodiments mentioned above, the gas flow passage cross-sectional area will be more concentrated to a central region between blades, which will change the global pressure distribution in an advantageous manner so as to reduce secondary flow losses, especially at the blade roots. In connection thereto it can be mentioned that GB2275308A discloses a gas turbine engine with a fan casing with straight members. However, since the straight members are connected to the outer ends of spokes extending inside the casing in the radial direction, the result is the opposite to that obtained with the second aspect of the invention, namely that the gas flow passage cross-sectional area will be more concentrated to the spokes.
In addition, the concentration of the passage cross-sectional area to a central region between blades will allow for the outer blade ends to be positioned closer to the axis of rotation of the engine. This will provide more space outside the blades, for example for oil conduits or air conduits, and it will also make assembly of the engine easier.
Preferably, two portions of the trace, at said cross-section, of the outer passage surface on either side of the respective blade, are parallel.